The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in several turbine stages which power the compressor, and produce additional power for powering an upstream fan in a turbofan aircraft application, or in driving an external shaft for a land or marine vehicle.
A typical turbine stage includes a turbine nozzle having a row of stator vanes which direct the combustion gases into a corresponding row of turbine rotor blades extending radially outwardly from a supporting rotor disk. A turbine shroud surrounds the rotor blades and provides a small clearance or gap with the blade tips for minimizing undesirable combustion gas leakage therepast.
The first stage turbine receives the hottest combustion gases from the combustor and requires cooling for ensuring a suitable useful life thereof. Cooling air is bled from the compressor and channeled through the hollow nozzle vanes and rotor blades for providing internal cooling thereof. Additional air is bled from the compressor and is channeled to the surrounding turbine shrouds for cooling thereof.
The prior art is crowded with various configurations for cooling the nozzle vanes, turbine blades, and turbine shrouds which vary in complexity and effectiveness. The amount of cooling air should be minimized for maximizing efficiency of the engine, yet sufficient air must be used for ensuring suitable component life.
Large gas turbine engines have correspondingly large vanes, blades, and shrouds which permit various forms of cooling configurations therein. However, small gas turbine engines have correspondingly smaller vanes, blades, and shrouds and therefore have limited space in which the cooling features may be incorporated, and correspondingly limit the types of cooling configurations which may be used.
For example, the turbine shrouds which surround the blade tips include conventional rails that mount in complementary supporting hooks in a hanger which limits the available space for introducing cooling features therein. The hanger itself includes rails which are mounted in complementary hooks in a hanger support, which support in turn is suitably mounted to a surrounding outer casing, such as the combustor case.
The nested configuration of the turbine shroud, supporting hanger, hanger support, and outer casing require suitable air circuits extending therethrough disposed in flow communication with the compressor for providing a portion of the compressor discharge pressure (CDP) air to cool the shrouds.
Shrouds themselves are typically formed in arcuate segments of a suitable high strength metal for withstanding the hot combustion gases, with the inner surface of the shroud typically being covered by a ceramic thermal barrier coating (TBC) joined to the shroud by an intervening metallic bond coat. The TBC provides effective thermal insulation for reducing the heat loads transmitted into the supporting shroud.
The shroud itself is typically cooled on its outer surface by the air bled from the compressor. Enhanced cooling of the shroud is typically provided by incorporating a thin sheet impingement baffle perforated with a pattern of small impingement holes. The baffle is suitably spaced outwardly of the shroud so that the cooling air is channeled through the individual impingement holes creating small jets of cooling air that impinge the back surface of the shroud for providing enhanced cooling thereof.
The cooling air is typically provided to the impingement baffle through corresponding inlet holes extending through the hanger either radially therethrough, or inclined therethrough with substantially axial orientation. In either configuration, a small number of large hanger inlets are provided around the circumference of the annular shroud support to feed the substantially larger number of small impingement holes found in the several segments of impingement baffles aligned circumferentially around the corresponding turbine shrouds.
In the large gas turbine engines, adequate space is typically available to discharge the large jets of cooling air through the hanger inlets with sufficient diffusion around the impingement baffles for reducing the velocity of the incoming air while increasing the static pressure thereof. In this way, a generally uniform static pressure distribution may be provided in the incoming cooling air to ensure substantial uniformity of impingement cooling through the multitude of impingement holes in the several impingement baffles.
However, in small gas turbine engines, or in large engines where space is limited, the configuration and orientation of the hanger inlets may be constrained and thereby limits the ability to adequately diffuse the cooling air prior to engagement with the impingement baffles.
Tests have been conducted in one type of small gas turbine engine in which the hanger inlets create corresponding jets of cooling air outside the impingement baffles with limited diffusion prior to passage through the impingement holes. The tests indicate that the high velocity jets of cooling air discharged from the hanger inlets can create local zones of relatively low static pressure, and correspondingly low flowrates of air through the local impingement holes. In this situation, the impingement holes within the direct local affects of the inlet jets are less effective for impingement cooling the backside of the turbine shrouds than those remote impingement holes offset laterally from the hanger inlets.
Accordingly, it is desired to provide an improved configuration for impingement cooling turbine shrouds notwithstanding the local jet flow from the hanger inlets.